United Technologies Corporation

United States of America As Represented by the Secretary of the Air Force

Raytheon Technologies Corporation

Eric J Ward

Edward B Thayer

James E Jones

Brian K Richardson

Kevin L Worley

Roger W Bursey

Philip R Scott

John H Maurer

Kenneth M Pesyna

Kenneth W Froemming

Jr Roger W Bursey

A variable vane assembly for a gas turbine engine includes at least one synchronization ring, a plurality of variable vanes connected to the synchronization ring, a plurality of torque boxes disposed circumferentially about the synchronization ring, a drive ring coupled to each of the torque boxes such that the drive ring causes an approximately identical torque to be applied to the synchronization ring from the torque boxes simultaneously, and an actuator coupled to one of the plurality of torque boxes and operable to drive the drive ring through the torque box.

Variable vane positioning apparatus for a gas turbine engine

A turbine frame for a gas turbine engine according to an example of the present disclosure includes a turbine case, a bearing support housing that defines a receptacle, and a tie rod preloaded in compression between the turbine case and the bearing support housing.

Turbine frame assembly for gas turbine engines

A gas turbine engine includes a first rotor shaft and a second rotor shaft around the first rotor shaft. A knife edge seal is radially between the first rotor shaft and the second rotor shaft, and is connected to one of the first rotor shaft or the second rotor shaft. The knife edge seal is sized radially to form a gap between the knife edge seal and the other of the first rotor shaft or the second rotor shaft. A plurality of blades having a general airfoil shape is between the first rotor shaft and the second rotor shaft proximate the knife edge seal. The plurality of blades is connected to the first rotor shaft or the second rotor shaft, and a second gap is formed radially between the plurality of blades and the other of the first rotor shaft or the second rotor shaft.

Intershaft bladelet seal

A compressor section for use in a gas turbine engine has a plurality of compressor stages. Each compressor stage is provided with a rotor hub, and each rotor hub mounts a plurality of compressor blades. A rotor drum extends between adjacent rotor hubs. A plurality of stator stages have an outer ring and a radially inner ring. A plurality of vanes extend between the outer and inner rings. An inner surface of the inner ring is spaced by a gap from an outer surface of the rotor drum. The rotor drum is provided with a plurality of blades having a general airfoil shape to resist leakage across the gap.

Gas turbine engine compressor stator seal

A rotational duct blocker for a gas turbine engine includes a duct blocker rotor and a duct blocker stator. The duct blocker rotor includes a plurality of first flow apertures extending axially therethrough, and a plurality of roller elements connected to an inner rotor platform. The first flow apertures are arranged circumferentially around the inner rotor platform. The duct blocker stator includes a plurality of second flow apertures extending axially therethrough, and a roller element track located radially within the inner rotor platform and in rolling contact with the roller elements. The second flow apertures are arranged circumferentially around an inner stator platform that is located axially adjacent to the inner rotor platform. The duct blocker rotor rotates relative to the duct blocker stator to regulate fluid flowing between the first flow apertures and the second flow apertures.

Gas turbine engine duct blocker that includes a duct blocker rotor with a plurality of roller elements

A mounting arrangement for a gas turbine engine blade having a root employs a lightweight, simple and economical leaf spring spacer which biases the blade root in a radially outward direction to minimize unwanted movement of the root within a conforming slot in a blade hub under conditions such as windmilling when centrifugal force alone is inadequate to tightly seal the root within the slot.

Gas turbine engine blade mounting arrangement

A valve assembly for supplying cooling air to a nozzle of a gas turbine engine. The nozzle can move between a first configuration and a second configuration. The valve assembly includes a first member having openings that receive cooling air. The valve assembly also includes a second member having openings. The second member resides adjacent the first member to receive the cooling air from the first member. Finally, the valve assembly includes actuator for moving the first member from a first position, which allows a first flow rate of cooling air to pass through the openings, to a second position, which allows a second flow rate of cooling air to pass through the openings. The second flow rate is less than the first flow rate.

Valve assembly for gas turbine engine

The flaps of the C/D exhaust nozzle for a gas turbine engine are arranged to obtain synchronous movement relative to each other and the interconnecting fulcrum links are attached to a coaxially mounted combined piston/sync ring by an H-shaped link. One link from one convergent flap is connected to one of the fulcrum links and the adjacent convergent flap is connected to the same fulcrum link. The redundant synchronization provided by the H-shaped links and flap connections provide damping of the flap train. The fulcrum links are judiciously mounted between adjacent convergent flaps.

C/D nozzle with synchronizing ring link suspension

A gap seal for convergent/divergent flap hinge ends to spaced disc and wall surfaces in a 2-D nozzle of a gas turbine engine is provided. The seal employs at least two cups resiliently mounted to the hinge ends, which cups independently translate axially and resiliently to contact the spaced surfaces at different or varying gaps to seal same. The translating cup seals of the invention thus maintain a constant seal in the gaps between the hinge ends and such surfaces at varying C/D hinge positions and varying operating conditions in the 2-D nozzle. That is, the cups independently maintain their seals across gaps as they expand or contract due to thermal changes as well as across gaps of different sizes.

Hinge end seal

The present invention comprises at least one active cooling apparatus incorporated into the nozzle liner of sidewall of an afterburner. The liner has a plurality of cooling air holes located under the apparatus. Within the liner is a supply of cooling air under high pressure. Over the holes in the liner is flexibly attached the active cooling apparatus which is composed of an inner manifold having a bottom layer with a plurality of receiving holes therein which are offset from the supply holes of the liner and a top layer also having a plurality of vent holes therein which are further offset from the holes in the bottom layer. Attached over the inner manifold is a hot cover layer which is located a small distance above the inner manifold top layer. The hot cover layer also has a plurality of flow path holes therein. A hot gas stream flowing over the hot cover layer bleeds into the inner manifold. Due to the differences in thermal expansion rates of the bottom layer, the top layer and the hot cover layer, the layers bow out from the liner. When this occurs, the cooling air can flow out into the active cooling apparatus and into the hot exhaust gas and in so doing reduces the temperature of the active cooling apparatus and allows it to bow back to the liner and close the holes therebetween once the hot stream stops.

Active cooling apparatus for afterburners

A thrust reverser in a two dimensional jet engine exhaust nozzle employs a plurality of reverser apparati that maintain a constant flow throat area, for example, in the reverse throat mode. A single reverser apparatus includes an actuator that operates through a lever upon linkage and a coupler to drive a set of parallel vanes and at least one independent vane. The parallel vanes are connected by a parallelogram of links to the coupler. The independent vanes move in a manner as defined by a cam having a preselected path therein that causes the independent vanes to move in a desired manner to provide the constant throat area.

Vane control mechanism

A flexible fairing 13 is provided which includes an aerodynamic surface 20 composed of a plurality of closely spaced, telescopic fairing members 14 and a moveable fairing panel 18. One end 15 of each fairing member 14 is pivotably secured to a divergent flap 11 of a thrust vectoring nozzle, and the other end 17 of the fairing member is pivotably secured to the fairing panel 18 which is slideably mounted to the static structure of a nozzle 19 of a gas turbine engine. The flexible fairing 13 adjusts to accommodate both translational and rotational movement of the divergent flap relative to an airframe fairing 10, thereby maintaining an aerodynamic surface 20 between the airframe fairing 10 and the divergent flap 11 of the nozzle.

Flexible fairing for airframe/nozzle interface

An external fairing flap (26) is provided with an hydraulic spring (42, 44) for causing the external flap (26) to follow the motion of an exhaust nozzle control flap (18). The external flap (26) is hinged (32) at the leading edge thereof, and contacts the control flap (18) via corresponding rubstrips (40, 36) at the flap trailing edges (28, 38).

Fairing flap arrangement

A convergent/divergent gas turbine engine exhaust nozzle having reverse thrust capability includes a concentric flap (22) and arc valve (36) pivotable about a common axis (34). A method and linkage (48) schedule the opening of the valve (36) in response to rotation of the flap (22) into a blocking position with respect to the normally aftward flow of the exhaust gas (12). The arc valve (36) is accelerated by the linkage (48) to full opening speed prior to unsealing an alternate, thrust reversing flow passage (38).

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West Palm Beach, FL, United States of America

Eric J Ward